All fighter aircraft projects are contractually required to perform a full-scale spin demonstration program. The purpose of the flight test program is to identify the types of spins that could be encountered inadvertently during future operational usage of the aircraft, and the control techniques that are required to return the aircraft to the normal flight regime. Other purposes for these programs could include the demonstration of aerodynamic configurations that are spin resistant, automatic spin avoidance and/or spin recovery techniques, etc. In any event, spin demonstration aircraft are required now and will be required in the future, and these aircraft must be equipped with an emergency recovery system that is guaranteed to terminate any unrecoverable spin mode that might be encountered.
One spin mode to which modern fighter type aircraft are susceptible is the flat spin, wherein the aircraft exhibit "spinning top" motions. This type of spin usually has a high rate of rotation, an angle of attack between 70.degree. and 90.degree. and effectively no spin radius, the aircraft spinning about an axis that passes through or near the center-of-gravity of the aircraft. To maintain a flat spin, the aircraft must balance the nose-down aerodynamic pitching moment with an opposing gyroscopic pitching moment. The magnitude of the aerodynamic pitching moment is a function of the aircraft configuration, dynamic pressure and angle of attack (usually increasing progressively up to 90.degree. angle of attack). The magnitude of the gyroscopic pitching moment is a function of the mass distribution and the product of the roll and yaw rates generated about the aircraft body axes. In a flat spin, the yaw rate is considerably greater than the roll rate. The yaw rate required for spin equilibrium is determined by the magnitude of the aerodynamic pitching moment and the aircraft mass distribution. The other requirement for spin equilibrium is that the aerodynamic yawing moment about the body yaw axis be zero (actually very slightly propelling, i.e., pro-spin) at this yaw rate. Obtaining a flat spin requires therefore, that a propelling aerodynamic yawing moment be generated at yaw rates below that required for balancing the aerodynamic pitching moment and that the magnitude of this yawing moment decrease (approaching a zero value) as the required yaw rate is attained. If a damping (anti-spin) yawing moment is generated below and at the required yaw rate the flat spin cannot be maintained.
For most aircraft configurations, the pro-spin yawing moment is produced by the forebody of the fuselage, the magnitude of the moment being a function of forebody geometry (i.e. length and type of cross section area), angle of attack and rate of rotation. Because an aircraft usually cannot recover from a developed flat spin through manipulation of the available aerodynamic controls, it is the spin which pilots fear most. It would obviously be desirable that this type of spin motion be made unattainable through use of some aerodynamic device, which would also facilitate recovery from this spin mode when encountered.
Emergency recovery systems used to date to generate an anti-spin yawing moment are complex, and usually incorporate a tail chute which is extremely inefficient when installed on modern aircraft that spin flat. In some instances, the chute size which is required for a particular type aircraft becomes impracticably large. In addition, the length of the riser line that attaches the tail chute to the aircraft is critical. If the riser line length is too short, the chute tends to collapse in the low dynamic pressure and reversed flow field that exists above the aircraft. If the riser line length is too long, the chute trails the aircraft at an angle which results in a nosedown pitching moment but no anti-spin yawing moment. Even the optimum riser line length results in a chute trail angle that contributes only a small anti-spin yawing moment. To compensate for the small anti-spin yawing moment, large parachutes are used. However, the use of large chutes results in off-design loads on the aircraft, which necessitates extensive internal and external reinforcement of the fuselage. In some cases, the reinforcement of the fuselage incurs changes in the mass distribution and external shape of the spin demonstration aircraft which jeopardizes the applicability of the results obtained from the testing. Some spin demonstration aircraft, therefore, require another device or arrangement to assist or replace the present emergency recovery tail chute system, and it would be desirable that this type of spin motion be made unattainable through use of an aerodynamic device in production aircraft.
The present invention provides a light weight and relatively inexpensive device or apparatus that can be easily installed on full-scale spin demonstration aircraft, yet is still highly effective in the recovery from a flat spin. The device can be armed either by the pilot of the aircraft, or can be automatically armed or controlled by a system that employs an air data computer which considers the yaw rate of the aircraft and/or the angle of attack of the aircraft. Once the device is armed, one embodiment automatically deploys when the nose section of the aircraft is subjected to the centrifugal force experienced in a spin. Thus, the present invention provides a spin recovery device which deploys passively (i.e., without an affirmative external action), and which is substantially completely reliable. An alternative arrangement or method could involve deployment by hydraulic actuators, for example, activated by an air data computer in response to spin forces or incipient spin conditions, or activated by the pilot.
In operation the invention converts the sign (direction) of the fuselage yawing moment from pro-spin (propelling) to anti-spin (damping) by changing the flow field over the fuselage forebody at very high angles of attack, this being accomplished by effectively changing the geometric characteristics of the forebody. In this manner the direction or magnitude of the side force developed on the forebody of the fuselage is modified, and consequently the overall yawing moment of the fuselage.
The spin control or recovery device comprises in combination with the high speed aircraft, spin control means generally located at the forebody of the aircraft for deploying, from a housed position, outwardly away from the normal contour of the nose of the aircraft in response to spin forces, such as the centrifugal force generated by the spin, or incipient spin conditions (spin imminent). In one preferred embodiment of the invention, the device comprises first and second outwardly convex clam shell flaps or doors which are pivotally mounted at the forward ends thereof on the nose of the aircraft in opposed symmetrical relationship about the longitudinal axis of the aircraft. Tether lines or other devices may be coupled to the doors to limit the outward angular movement of the doors to an optimum value upon deployment thereof, and means for locking the doors in the housed position may be provided.
Other features and advantages of the invention will be set forth in or apparent from the following detailed description of a presently preferred embodiment, taken in conjunction with the appended drawings.